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Why doesn't temperature affect Pressure Altitude?
In this picture:

it states the tops and bottoms of the 3 columns of air have the same pressure.
This explains the error in altitude indications given in a non-standard temperature.
Pressure only drops 1 inHg per 1000 ft. when temperature is standard.
Assuming that the bottoms of the column of air are at sea level, why aren't pressure altitude readings (setting altimeter to 29.92) affected by the True altitude formula, TA = IA + (4ft. x (1000s of ft) x (OAT - ISA))?
I was always told temperature does not affect pressure altitude readings because the atmosphere is not a fixed volume of air. If temperature increases, it will also expand in volume, like how the picture explains.
I could see how pressure altitude wouldn't be affected by 4ft. per 1000ft per (delta°C ISA) only if the picture assumed the bottoms of the pressure started at ground level.
Or does it affect pressure altitude, just very slightly and non-significant? Not as much as if you were in a fixed volume of air...

What you read on the altimeter is not a "pressure altitude" but the "indicated altitude". Indicated altitude only depends on outside air pressure not outside air temperature. That's why the three airplanes in the picture read the same altitude, because the outside pressure is the same.

The picture you have displayed is correct. But, your understanding of what it is saying is incorrect. The picture is explaining the reason your Indicated Altitude will change with temperature for the same True Altitude. And, your True Altitude will change with temperature (a potentially dangerous situation) for the same given Indicated Altitude.
Your misunderstanding may stem from the fact that the Indicated Altitudes and the field level barometric pressure (or at least the Altimeter Kollsman Window) setting are the same in all three temperature scenarios. This is not the case for the pressure at 10,000 feet MSL True Altitude. The atmospheric static pressure at altitude for the three scenarios is not given. And, they can not be equal to 29.92.
The static atmospheric pressure is higher for the cold scenario and lower for the warm scenario at 100,000 feet MSL True Altitude. You would think that this would be the opposite way around. And, it is for air confined in a container. But, not for unconfined air or air confined in different sized containers.
Atmospheric air will become less dense and rise when warmed. This will create a partial void of air below it. This partial void is an area of low pressure since the surrounding air has yet to move in to fully replace the vacating air. If the air were to be cooled, it would sink onto the air below it. This would create an area of high pressure as the existing air has not moved outward in sufficient quantity yet.
How does this affect the different types of Altitude? Like this:

True Altitude is the height above MSL measured in actual feet of distance.
Pressure is just a measurement of the amount of air interacting with the measurement device and the average amount of energy each air particle has. Temperature will affect the amount of energy possessed by each air particle proportionately. This will either cause the  pressure to change proportionately. Or, it will cause the volume and density of the air parcel to change proportionately.
Barometric Pressure (Altimeter Setting, Kollsman Window Setting) is just the static atmospheric pressure at field elevation adjusted to show what the pressure should be at Mean Sea Level.
Indicated Altitude is just the reading you receive from your Altimeter. Your altimeter is just a very accurate pressure gauge (a barometer) calibrated to show the height above a specific reference line of pressure. The pressure and density of air at lower altitudes are greater than they are at higher altitudes. This is due to the sheer weight of the air from above pushing down and compressing the air below. The amounts of changes of air pressure and air density are known. This is based on observationally, scientifically and mathematically derived formulae and estimations that are based on the International Standard Atmosphere. On an ISA day, that reference line is equal to the Mean Sea Level. That is not always convenient since the atmosphere is not always at standard pressure and temperature at Mean Sea Level. So, most altimeters can be adjusted to account for non-standard barometric pressure.
Pressure Altitude is just the height above the mathematically derived line representing where the International Standard Atmosphere pressure of 29.92 inches of Mercury should be felt. Pressure Altitude is the height in feet you would feel the same pressure on a non-standard day as you would on an ISA standard day in feet MSL.  If the atmospheric pressure at MSL (or the barometric pressure at field elevation) is not standard, the two beginning levels will not be in the same place. You would have to make an adjustment to the height of MSL to derive the Pressure Altitude.   The pressure of air is derived from its volume and energy content. Temperature is a factor in this only if the air is confined or in a container.
Density Altitude is just the height above the mathematically derived line representing where the International Standard Atmosphere density. Density Altitude is the height in MSL where you would feel the same density on a non-standard day as you would on an ISA standard day.   If the atmospheric density at MSL is not standard, the three beginning levels will not be in the same place. You would have to make an adjustment to the height of MSL to derive the Density Altitude.   The density of air is derived from its weight and volume. Both pressure and temperature are factors in this only if the air is unconfined or in a container.

You derive Density Altitude by adjusting Pressure Altitude by atmospheric temperature. You derive Pressure Altitude by adjusting the True Altitude by the atmospheric pressure. You have to have each to get the next.
Case in point. If the air were confined to nice, neat, little columns with defined and rigid lateral boundaries and adjustable vertical boundaries, the average pressure in each container would remain the same. The heights of each column would be different based on temperature. For extremely tall columns, the pressures at the bottoms would all be equal. And, the pressures at the top would all be equal. But, the rate at which the pressures would change with height would be different. Since Indicated Altitude is just another way to represent the static atmospheric pressure at altitude, the Indicated Altitude of 10,000 feet MSL would be at different True Altitudes based on temperature.

In your 3rd bullet, you say that your altimeter setting is static pressure at field elevation adjusted to show the pressure at 0 MSL. If the airport is 1000 ft. above MSL, then the actual pressure on ground should be lower than the kollsman window right? For example, in the picture provided, if they were all set to 29.92, why would being at the field (1000ft) give the correct pressure altitude without being affected by temperature compared to being at 10,000 ft. MSL? My understanding of the gradients being stretched out is what causes the 4ft per degree per 1000ft.

The easiest way to explain my question is say the airport was a quarter way up in altitude on the middle column in the picture.  Why isn't it affected by temperature when setting it to 29.92 even though the actual pressure would be slightly lower than MSL?

@user7828137 - I can see how my wording is a little confusing. The barometric pressure that is used to set your altimeter Kollsman window would be the barometric pressure at Mean Sea Level. If you are at an airfield whose field elevation is above Mean Sea Level, there is no way to measure the actual barometric pressure below ground level. The static atmospheric pressure at field elevation is measured. Then it is adjusted for the higher than sea level elevation. The static pressure measured will be lower than the barometric pressure. But the adjustment is fairly predictable.

@user7828137 - If you set your Altimeter to the correct barometric pressure at MSL, it should show you your correct field elevation to within 75 feet. Even though it is actually reading a lower static atmospheric pressure. If you were set your Altimeter to the static pressure measured at field elevation by a non-calibrated barometer, your altimeter would read zero feet. If you were to set your altimeter to zero feet, your Kollsman window would read a lower setting than what the current setting would be for that field if the field were above MSL.

The picture you have given does not represent three airports at different field elevations. It represents the same airport or three airports at the same field elevation. True Altitude references MSL. Absolute Altitude references AGL. The actual field elevation is irrelevant in this picture. Your idea on temperature affecting pressure is correct if the air were in a rigid container. Instead, the air is unbounded. Think of the columns as cylinders with movable pistons. So, instead of the pressure changing with temperature. The volume changes. And, weight stays the same.

Okay, and because the measurement was taken at field elevation, this is why there would be no error in pressure altitude no matter what the temperature is? If they actually took the pressure at MSL and gave the airport that number, THEN would there be error in our indicated altitude based on extreme temperatures on the ground at the airport, say 3000ft?

The bottom of these 3 columns of air are at 0 MSL, correct? Not the airport? In reference to your last comment.

@user7828137 - You are way overthinking this. The barometric pressure is what the static pressure would be if there was air exposed to the atmosphere at MSL. It does not take into consideration the pressure of the much denser ground between MSL and field elevation. It’s like there was an open well dug down to MSL at the site of measurement. The bottom of the 3 columns are at MSL regardless of field elevation. And, 3000 feet in either direction would not produce extreme temperatures. Temperature will affect the density of the air.

Because Pressure altitude is a measure of the weight of the air above you. Temperature does not affect that. Heating the air just makes it expand. It still weighs the same.

I'm confused on when temperature affects the pressure vs. When it doesn't, because it does change 4 ft per per degree from standard temperature per 1000 ft when flying. I understand  (I think) that at 0 MSL, the weight above is the same, but for an airport at 3000 ft., why does it not change there? You are already 3000 ft up in the pressure gradient so isn't there room for error in those 3000 ft.?

@user7828137 - The altimeter setting is not the static atmospheric pressure at field elevation. It is what the static atmospheric pressure would be if the field were at Mean Sea Level. If it were not consistently given referencing MSL, aircraft Indicated Altitudes would show AGL instead. Two different aircraft flying in the same general area with 1000 foot separation and the same altimeter setting would be in conflict if one were flying above an area with 1000 feet of difference in field elevation.

@user7828137 - From your comment, I can see that you are putting the cart before the horse. It is not the temperature that is affecting the pressure (not directly at least). It is the pressure that is affecting the temperature. If the atmosphere has a standard lapse rate, the temperature of the air will decrease by 2°C for every 1000 feet of altitude increase. This is due to adiabatic cooling (cooling due to decreasing pressure).

First of all, why would anything in the 3000 feet below you affect the pressure? That air cannot be pressing on you. It is below you. You can stand on a 20 ton block of concrete without ill effects... But put the block on your head ... The formula you reference is a part of a shortcut used to calculate pressure altitude when you know the local altimeter setting. It assumes that the altimeter setting is the measure of the air that would be above you if you were at Sea Level at the same place. You cannot draw any inference from this as to how pressure altitude changes from temperature.

@CharlesBretana "First of all, why would anything in the 3000 feet below you affect the pressure?" In my opinion, it would. The amount of air above you is determined by the amount of air below you; the total mass of air is fixed, therefore the mass above is what remains after subtracting the mass below from the total mass. If the temperature was to fall, density will increase, resulting in more air below. Shouldn't that decrease the weight of the air above?

@Aditya, no actually it doesn't. This is, actually, a core concept used in hydraulic engineering and in fluid mechanics. The relationship between indicated altitude and pressure altitude may (does) change, but not because the total mass above you remains constant.   The pressure you feel at any altitude is not based on a subtraction process. It is created by the force of the air molecules pushing on you from the weight of those above.

If you heat up a local section of air, and remain at the same absolute altitude,  then yes, some of the air below you may, as it expands, move above you. But that is not guaranteed. It might move sideways. That's why we have wind, and storms.

The weight of the air above you is very much affected by temperature!!

As already stated, pressure is effectively the weight of the air above the aircraft.  If you were (for example) to double the absolute temperature of the entire atmosphere (from say 300K at sea level to 600K then the atmosphere would double in volume, assuming it’s an ideal gas, which isn’t a million miles from the truth, and pressure altitudes would increase by almost a factor of two (but not quite, with the Earth being spherical).  However, a localised temperature change would tend to result in expanding air spreading out into the surrounding airmass rather than generating a localised lump in the upper atmosphere.

To me the short answer to "Why doesn't temperature affect pressure altitude?" is "Well, it does!!!"
I know where this confusion comes from. Most, if not all, flight instructors repeat this mantra backed by the fact that temperature does not show up in the formula to compute the pressure altitude. Or other nonsense like "Because Pressure altitude is a measure of the weight of the air above you. Temperature does not affect that"
Yet, I wonder how many people, if asked "is atmospheric pressure affected by temperature?", would reply "No, it's not". Everybody knows that temperature affects atmospheric pressure, right?
But that's basically the same question as "Does temperature affect pressure altitude?".
That's because:

Pressure altitude is just a fancy way to measure atmospheric pressure
Altimeters simply measure atmospheric (static) pressure and display it to you as an altitude (see the math here)
Altimeters with QNH set to 29.92 measure pressure altitude
Changes of temperature cause change of air pressure

EXAMPLE: Suppose you're on top of a very high building. If you have an airplane altimeter with altimeter setting set to 29.92, it will show 2 different values (i.e. pressure altitudes) on cold days and hot days. But the height of the building (that represents real altitude in this analogy) doesn't change.
So if temperature affects air pressure, then it affects the reading of the altimeter, then it affects the value of pressure altitude.

By definition temperature is not factored in to calculation of pressure altitude. Pressure altitude is always the indicated barometric altitude with 29.92 used as a barometric offset. Temperature will affect the altitude of aircraft flying at pressure altitude though.

If you were at an airport on two different days at 5000 MSL with the exact same barometric pressure say 30.92, but one day was freezing cold and the other day was very hot, would you have two different "pressure altitudes" then?

@user7828137 - No. You would have two different density altitudes with the same pressure altitude. That is why density altitude is taken into consideration when calculating aircraft performance. Pressure altitude is only a way of deriving density altitude. Pressure altitude can be easily measure by setting your altimeter to 29.92 inches of Mercury. Pressure altitude is the altitude at which your ears think they are it will affect the pressure gradient of your body. Density Altitude affects the oxygen absorption of your cells and the performance of your engine.

@DeanF. great explanation. There is only one thing that I dont quite connect. how come the barometric pressure at this airport of 30.92 hasnt changed for a freezing cold vs a hot day? I found online that the factors to change the barometric pressure is moisture, temperature and elevation or msl. Thus, why is it that temperature is not a factor to the change of barometric pressure? because if it is then you would have to different readings, two different Pressure altitudes, correct? Thank you for the clarification

Pressure Altitude at a specific point is based on a few different variables. One of which is the conditions of the surrounding areas. Even if the temperature changes, the barometric pressure can stay the same. This is primarily due to their not being volume restrictions on the air. If the air were in a sealed container, this would not be the case. A change in temperature in a sealed container causes a change in pressure.


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Why did TWA800 not stall?
I believe the NTSB account, but based on this graph, http://twa800.com/lahr/zoom-climb.htm, why did the flight not stall after the nose separated?

Did you read the report? https://twa800.com/report/finalreport.pdf specifically from page 94

In a nutshell, the aircraft had enough energy to climb without exceeding its stall AOA.  The expert stated the plane should have "immediately stalled", but the incredible amounts of mass and velocity associated with a plane that large demand a closer analysis to the physics.
When the nose breaks free, lift and weight are out of balance. When an aerobatic plane applies full up elevator, lift and weight are out of balance.
There are two "bookends" to what happens next.  The upper is G limits are exceeded and the wings tear off.  The lower is stall AOA is exceeded.  In between is an attempted loop.
Apparently, the rugged 747 had enough velocity to climb, sans nose, for a while.  Really a testament to its design.
Even after stalling, its momentum would have carried still higher, so it is difficult to determine when exactly it stalled.
Factors which determine "stallability" include wing loading, wing configuration, and tail volume. A very high wing loading with a very thin wing (as seen on many supersonic fighter aircraft) will stall much faster.  Deploying slats and partial flaps greatly helps a plane tolerate higher angles of attack.  Large tail volume also increasingly inhibits higher and higher AOA.
Essentially, it is a race between change in flight path and change in AOA within the G limits of the aircraft.

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Why are some modern military trainer aircraft so bulky?
Modern military trainer aircraft are a lot wider and bulkier than the previous generation. Especially the canopy looks oversized.

You can see this in the Boeing/Saab T-7 Red Hawk, Alenia Aermacchi M-346 Master (developed from the Yakovlev Yak-130) and Hongdu JL-10.
Development for these aircraft was started in the late 90s / early 2000s. All of these aircraft have a secondary light combat role. They are supersonic capable.
The seating position is staggered, probably so that both pilots (or student / instructor) have good visibility. This might explain the enlarged canopy. The wider cockpit is likely more comfortable and enables longer missions as well.
Additional avionics and increased fuel capacity may explain the overall increase in volume.
There seem to be some underlying reasons for these designs, but they are unclear to me.
What prompted these changes in the designs?

The helmets have grown bigger with helmet-mounted targeting, night vision and all that equipment which can be attached to a modern helmet. That requires more space, hence the canopies have become wider and higher.

I wouldn't say the current gen trainers are that bulky (compare the T-38 to its subsonic contemporary, the T-37 Tweet), more that the T-38's supersonic capabilities forced them to be slender)

You mentioned the oversized canopy for two, staggered for better visibility, and really that is the main feature that stands out.  Consider that trainers are typically quite a bit smaller than their full combat brethren and this feature is even more out of proportion.  Also, full sized combat 2 seaters generally have a Weapon System Officer in the back seat, so good forward visibility isn't as crucial as it is in a pilot trainer so less upward stagger of the back seat is needed.
The human body doesn't scale up and down like an airplane design, it is the constant that must be designed around. Think about it, you are fitting two 6 foot tall people in a fairly small plane vs often one in a larger plane; naturally the space for the bodies will be a much larger overall percentage of the size of the  trainer.
If you were to compare a 2 seat trainer next to a larger 2 seat combat aircraft like the F14 Tomcat, (with its huge engines...) I think you would find that the cockpit areas are actually pretty similar in size.  They are just more pronounced in the smaller plane.
Other than the cockpit area they really aren't "bulky" at all.  In fact the two pictured are VERY compact and streamlined, by just about any standard.

Also the F404 is larger than the two J85 engines in the T-38.

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Could the Perlan II glider be used for nearspace tourism?
I asked the same question about the SR-71 Blackbird, and I'm also curious about the Perlan II sailplane which travelled above 70,000 ft, and is planned to be flown to 90,000 ft, if it could be used for nearspace tourism and what some issues might be with that. Its starting location would have to be near a high enough mountain range where it can use enough lift to climb that high, and this is best at certain times of the year I guess. A major advantage would be that this form of nearspace tourism would be 100% environmentally friendly (for those who bother).

Well, environmentally friendly not counting the carbon footprint of launching, and of airline flights and ground transport to get the tourist to the launch site, and then back home again.

Considering that space is at 330,000ft I'm not sure if I'd call 90,000ft near-space.

Updrafts which would carry a glider to 70,000 ft are extremely rare. How long had the Perlan II crew to wait for the right conditions? This would make the experience extremely boring for tourists and impossible to plan ahead.

@GdD I wrote nearspace without dash, one might call the area between 60,000 ft (Armstrong limit) and 282,000 ft (Mesopause) nearspace or mesospace. Space begins at the mesopause where there are also Kármán stall speeds depending on aircraft.

@ZeissIkon Maybe tourists came by sailship and e-mobile, and the glider is being launched on a winch or from an electrically-powered carrier plane.

@PeterKämpf The Andean Wave (which the Perlan gliders have used) is a pretty dependable phenomenon.

@Giovanni And maybe I'll win $200 million in the lottery...

@ZeissIkon So you oppose (near)space tourism per se?

What exactly are you defining as "nearspace tourism"? It's a very different question if you define it as >60k feet vs. something higher than the Perlan II has achieved. If it's the first, then we know the Perlan II can achieve it, and the question is asking about challenges with using this for tourism, but if its a higher threshold, then we get into technical questions about how high a similar glider can go. Which one is it?

No, I oppose unfounded optimism, here seen in the form of claiming nearspace tourism can be eco-friendly.

Also, unrelated, but this question and your SR-71 question are from separate accounts with the same name. You should talk to the mods about getting them merged

@ZeissIkon The flight itself is (also the carrier flight if it's carried by an e-plane to altitude). If this isn't eco-friendly for you, I don't know what is.

@divibisan I'm asking about the challenges with using the Perlan II for mesospace tourism (to the altitude it can achieve, ~77,000 ft is its record).

No, Perlan II can't be used for tourism.  It has only two seats, and every description of it talks about the crew of two, or the two pilots, or the chief pilot and co-pilot.  Someone who gets certified for the type and carries that much responsibility isn't called a tourist.


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Is it dangerous for a supersonic aircraft to fly at exactly Mach 1?
If a plane that is designed for supersonic flight, say the Concorde, kept flying at exactly the speed of sound, would there be any danger in that? If so, what could get dangerous at constant Mach 1 flight and how long can a plane survive in flying at Mach 1?

Why exactly is Mach 1 supposed to be especially dangerous?

@Abdullah I ain't sure, this is why I ask. I think you're flying more smoothly much below or much above Mach 1, but when flying at Mach 1 your plane's engines produce a sound you are flying at the same speed of it with it. This sounds dangerous to me.

Take a look at the Boeing Sonic Cruiser, which was intended to fly in the transsonic region for its cruise period.

How could that be?

Make sure to read this answer to understand what is special about Mach 1.
Pilots regularly report that, approaching Mach 1, the airplane is shaken by oscillating shocks (they don't use that term, though, they speak of buffeting) and once the Mach meter needle crosses the 1, the airplane becomes calm and flight smooth again. Since the flow speed around the airplane is normally a bit higher than flight speed due to the displacement effect and the need to produce lift, supersonic pockets start to occur as early as at Mach 0.6 (depending on the particular design) and expand the closer flight speed comes to Mach 1.
Swept wings delay those Mach effects and the peak of the drag coefficient of supersonic-capable designs is also delayed to somewhere between Mach 1 and 1.2. While the drag coefficient normally drops a bit past this peak, drag itself normally rises with speed. Flight at Mach 1 is particularly inefficient but, apart from that, drag is not an issue.
What is more important is the shift in the center of pressure when changing from sub- to supersonic flight. It will move aft and requires the pilot to re-trim the airplane. Supersonic-capable designs normally use full-flying tails for that and delta airplanes raise their trailing edge flaps. Concorde added pumping fuel aft and the XB-70 reduced the rear part of the wing by folding its wing tips down. When flying at Mach 1, the airplane is smack in the middle of the largest pitch moment change over speed range, so any change in speed means also a strong change in pitch moment. When no trim change accompanies that change in speed, the airplane will be unstable in pitch: Flying faster will move the center of pressure aft and make the airplane dive (Mach tuck), speeding up more, and vice versa. Reducing speed might even lead to structural failure from the sudden pitch-up and increase in load factor if not corrected by control deflections.
That is what makes flight at Mach 1 unadvisable. If the pilot or the FCS keep pitch under control, flight at Mach 1 can continue until the fuel runs out and the airplane hits the ground, causing speed to drop dramatically.

@ymb1 Being particularly slender, the SR-71 was uniquely unaffected by Mach effects. Also, its shape reduced the pitch trim change significantly. That the air data system showed jumps in Mach stems from different techniques to determine Mach sub- and supersonically (see Wikipedia).

what about buffeting and flutter and shock induced stall.

@Abdullah Flutter can occur at almost any speed. Shock-induced stall is a problem for subsonic designs which fly near the coffin corner. While both are hazardous in special conditions, they are not bound to Mach 1.

Drag forces grow quickly as you approach the speed of sound, and then fall off somewhat after you are above the speed of sound. This means flying near the speed of sound puts you in a high-drag regime, where you are putting work into the flow as shock waves are trying to establish themselves at various points on the airframe. Unless your airframe is perfectly symmetric and smooth throughout, those shock waves can't be expected to be perfectly symmetric either, and the result will likely be quite a bit of buffeting and banging around- and really excessive fuel burn.
That's going to compromise the controllability of the plane as well as expose it to a barrage of shock loads that will eat up the fatigue lifetime of the structure, and furnish a rough ride for your passengers.

Buffeting is gone at Mach 1. It only occurs when flight speed is still subsonic so shocks sit somewhere along the chord and can move chordwise instead of fixed at the trailing edge which happens at and above Mach 1.

@PeterKämpf, thanks for posting your own answer to this, it is much more thorough than mine. -NN

Mach buffet precedes mach, and begins in the transonic range.  The transonic range does not begin or end at 1.0 M1.  It begins typically around .85 and may continue to 1.2 to 1.5.  Buffeting may occur at any point within that range.
Mach is not a point that is reached uniformly by the entire aircraft.  This is to say, mach airflow will be reached at some points of the aircraft sooner than others.  An aircraft may be at the lower end of the transonic range, while some portion of the aircraft is experiencing mach airflow.  The 747, for example, experiences mach airflow at the rear of the upper deck supernal "hump" or dome.  On freight 747's with little insulation, the airflow at the rear of the upper deck may be heard banging and snapping (this is not a "supersonic pop" or "bang," but rather airflow disruption and interference with surrounding airflow; it is a place where local airflow exceeds M1 in transonic flight).
In a sense, mach buffet can be thought of as the way a motorcycle feels on the highway, driving behind a big rig truck, or approaching the front of the truck, in the adjacent lane, on the highway.  That shaking and buffeting from disturbed air.  Get ahead of the truck, the buffeting goes away.  Approaching the front of the truck, one begins to encounter air deflected from the front of the truck, and buffeting.  Get just past it, and it smoothes out.
A big difference is that at highway speeds, the air isn't considered compressible.  In flight, approaching mach, compressibility and shock waves become a factor; more than one occurs, both ahead of the wing and along the cord.  When the shock waves have passed along the cord, the drag rise associated with approaching mach, and aft movement of the shock wave, are no longer a factor and the aircraft operates more smoothly.  Other effects such as mach tuck, also associated with movement of the center of pressure stabilize.
The buffeting and mach effects are most prevalent in the transonic region.


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Why am I getting a low L/D of 4.3 for this supercritical airfoil in Autodesk CFD?
I am a high school student learning to analyze aircraft models in Autodesk CFD for the first time. My current project is analyzing a supercritical airfoil on Autodesk CFD (that was designed in Autodesk Fusion360).
Can I get help with the following : I researched several turbulence models in Autodesk and I chose the SST k omega turbulence model, with a 250 m/s inlet velocity, 0 gauge pressure , unknown outlet and a compressible flow.
I am getting a lift to drag ratio of approximately 4.3. I think the lift to drag ratio is too low. I am suspecting the analysis conditions are to blame. Are these the correct input parameters? Am I missing anything else?
Here is a screenshot of the analysis: I ran the airfoil analysis at angle of attack of 5°.


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What is the feasibility of fitting all aircraft with black boxes?
In the wake of the tragic helicopter crash on January 26,2020 which killed Kobe Bryant and eight other people, it would seem to make sense to follow transport category aircraft standards and fit all aircraft, except ultralights, LTA unpowered, or powered parachute, with a cockpit voice recorder and a flight data and GPS telemetry recorder?
These systems aren’t that heavy - roughly 10 lbs or so - so it would represent a minimal loss in useful load compared with other standard of optional equipment.  I’m guessing with a little ingenuity, such systems can be made lighter and better tailored to light aircraft.  And these systems have been indispensable in accident investigations and may well greatly improve light aircraft safety.  Modern AHARS systems retain some telemetry data, one of which was responsible for piecing together this stall spin accident of a Cirrus.

One downside may be the requirement to place these systems in the tail boom of an aircraft for maximum survivability in an accident, and placing even light equipment that far from the aircraft datum, can cause dramatic shifts in Moment and CG location.

They’re expensive.

Lol.  Name one thing in aviation that ISN’T expensive!

Somewhat.  But regulations alone don’t seems satisfactory as laws can be changed.

Are you going to tell every GA owner that his $70K airplane needs to have a $25K flight data recorder installed, assuming the rest of his aircraft can actually provide that data? Look how much of a pain the ADS-B mandate was

Just because they don't have an "all encompassing" black box, doesn't mean they don't have the basic equivalent of one. For example most GPS maps record data in the background, as well as things like ADS-B units. If these units survive enough to have the data extracted, they are used to build a "black box" to help investigators.

There are serious privacy concerns:  Why should a private pilot, flying himself and his close buddies, be mandated to have private conversations recorded?  What if he's discussing business dealings, or extra-marital affairs and a court tries to subpoena the black-box for something not related to aviation safety at all!

@abelenky can/does that happen with commercial flight black boxen? Granted, they only hold but so much info and a commercial plane is used many times more than a private, thus overwriting last week's conversations, but, as a serious question, is that a concern?

That sort of privacy issue reared its ugly head in the circumstance of company owned EFBs tracking the whereabouts of the employees of one of the regional airlines.

@abelenky, the information is not made publicly available and could only be released to law enforcement or other authorities under a court ordered search warrant.  Keep in mind similar risks exist by simply carrying a cell phone every where you go.

A commercial plane is owned and operated by the air carrier, under a federal charter.   A private plane is very different.  It is unsettled and largely unexplored legal territory.  But it would be problematic to mandate voice recorders on someone's private two-seater puddle-hopper.

@Ron Beyer: I serious doubt that most light aircraft have a GPS, and of those that do, most probably would not survive a crash.  Heck, I've seen phones & tablets break just from being dropped on the floor.

@jamesqf Every small aircraft I've been in has one, even if it's a portable unit like an Aera. The device itself doesn't need to survive, data can be retrieved from bare chips.

@jamesqf, they're becoming more and more ubiquitous.  If your airplane has an IFD, it will have an integral GNSS system or GNSS equipped panel mounted units (eg, IFD440/540, etc.).  More and more aircraft are receiving GNSS retrofits as part of the ADS-B out requirement starting as this year.

It would be very difficult to mandate new equipment into older aircraft. Look how much it took to mandate ADS-b. Mandating it into new or future built aircraft would be easier. We may try to incorporate the telemetry-like data we are getting now into some type of solution. Companies like Sentry and Stratus could design a portable unit that records all AHRS, GPS, traffic, cabin pressure altitude, and carbon monoxide presence data for less than $1000. Combine that with a personal electronic device like a tablet or smartphone and a patch cable or Bluetooth connector, and you have most of the capabilities of an FDR/CVR combo. It would not record the position of flight and engine controls, or engine performance data, however.
The use of this may be easier to mandate. $1000 is fairly reasonable in aviation economics.
Making it mandatory for ALL passenger and property carrying commercial flights may be more feasible and likely. The federal government would implement it under the auspices of protection of the consumer or for the good of the public.

It should be mandatory for all celebrity-carrying flights also.

@rclocher3: Why?  I mean, what would the most sophisticated black boxes tell us, beyond the obvious (and unfortunately too common) flying lower than the tops of the hills in IMC conditions?

I was kidding @jamesqf.


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Why do X-wing configurations missile/guide bombs use symmetric airfoils?
From all the missiles/guide bombs I have come across, I've always observed that a symmetric airfoil is used for both the wing and tail.

Why not have a cambered airfoil for the wing and a symmetric airfoil for the tail? It would provide a better range for the missile/guide bomb.

If an airfoil were to be selected for X-tail missile/guide bomb configurations, what is the selection criteria? I'm trying to better understand the reason for this generally accepted design decision.

Any answers/suggestions would be appreciated.

Assymetrical wings on missiles/guided bombs do exist. However the goal of all missiles/guided bombs is to hit or come close to its target, which mean different things for various missiles. This drives their design:

Symmetrical airfoils on missiles are used in order to guarantee uniform turning capability. This enables the missile to immediately turn in either direction, therefore not losing time to first roll the entire missile before turning, therefore making the missile more maneuvrable. This is especially needed for air-to-air missiles like the R73-Wympel or the Diel IRIS-T. Cruise Missiles, on the other hand, can be less maneuverable (ground targets seldomly move), which is why they do not necessarily have symmetrical wings. For example, if you look closely at the image of the Tomahawk missile, you will notice the unsymmetrical wingshape.  Such missiles use the bank-to-turn control strategy. *.

I do not know too much about this, but the main design goals for these again are good turn capabilities (aka, pull a lot of gs), and supersonic performance. Therefore basically the same considerations you have as for any other airfoil with the exception that the requirements are a bit more extreme.

* An interesting mix is the MBDA Meteor. It uses bank-to-turn because it has underbelly intakes. I cannot find the source, but apparently in the endgame the control mode is switched to skid-to-turn as it is more important to hit the target, then to keep the motor running

Most missiles fly at such high speeds and manoeuvre so violently that "up" means very little to them, it is simply a mild error to be corrected for.
Cruciform wings and tails are designed solely for manoeuvrability in whatever direction the missile needs at that moment. After a bit of that, there is no predicting which way up the missile will be by then anyway. A symmetrical solution gives optimum performance.
Typically, if both are present, an X-tail will be set at 45 deg to the X-wing, to minimise aerodynamic interference.

It's not guaranteed that the missile/bomb will stay upright relative to an asymmetrical wing throughout its whole flight, and if it were to wind up inverted, the range would be shorter and the package would fall short of its target. That might mean on civilians or friendly forces, and that's not good. Asymmetrical airfoils are effective, and they get the job done- it's simpler that way.


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What are the random, differently-coloured rectangular patches on the top of this Boeing 737's fuselage?
The top section of this green-painted Boeing 737's fuselage was slightly faded and I noticed a few dark-green, rectangular patches on its top. They look like patch kits on a bicycle tyre's inner tube. What are they? My understanding is that hairline cracks/weaker spots in the fuselage can't simply be patched like this?

Cracks and weak spots are simply patched (after you cut out the cracked/weakened area to keep it from spawning fatigue cracks).

They aren't "repair patches" themselves, they are repainted areas where some sort inspection and/or repair was done to the underlying structure, that required removal of the paint in the area. What the inspection or repairs were done?  Well, you'd have to have been there to know, but the 737 does have a lot of structural inspection requirements related to fatigue cracking in fuselage frames and skins.
There may be actually be skin outside doubler patches riveted on in those dark spots to reinforce areas where crack repairs were done, but the borders of the dark spots are just paint masking lines and it's hard to tell in the photo what was done.
Another possibility is that surface corrosion was found at those spots and the paint in the area was stripped, the corrosion was blended out and the zone repainted.
They should probably have their paint guy do some mixing to match the fresh paint to the faded paint, and apply the paint with an edge feathering technique to avoid the sharp masking lines (you have the edges of the mask supported above the surface to allow some paint to spread below, to create a soft edge), so that folks in gate areas looking down at their incoming flights would be less alarmed by the Quilting Bee look.

I think it is better to paint with the fresh color of the same area and wait it to fade and blend in instead of matching the faded color and the matched color fading in its own fashion.

This is getting more into the realm of DIY.SE, but perfectly matching a patch of paint, particularly on an exterior surface that gets long exposure to the elements, can be next to impossible.  Even if you get it looking right immediately after painting, the fresh paint and old paint may fade at different rates and the difference can actually increase over time.  It'd almost be easier to repaint the whole plane.

Oh, one other point.  This color discrepancy is much more obvious when using bright, vibrant colors like the green in this example.  This might be why you don't see this very often, because most planes are painted predominantly white (possibly partly for this reason).  The paint can't noticeably fade as much when it was already white to begin with.


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Light aircraft elevator/aileron/direction not working
What is the correct procedure if a light aircraft (under1000kg) elevator/aileron/direction stop working? Is it alright to use the ballistic parachute(if the aircraft is equipped with one)?
If not, then in what situation ballistic parachute is used?

Anything that gets you on the ground safely in an emergency is an acceptable solution.

Welcome to aviation.SE! You seem to have two separate questions here and it would probably be best to ask them separately. The [tour] may be helpful to see how the site works.

Related, although it's about a 737

If a major flight control surface fails in flight, you'll be lucky if the ballistic parachute can save you.  An aileron, for instance, that goes "hard over", will put even the most docile training aircraft into a tight spiral or a roll that will quickly become a spin (because the airplane doesn't have enough speed for a complete roll and will stall).  A stuck elevator could result in anything from a stall to a steep dive.
In this situation, immediate deployment of a ballistic parachute might prevent fatalities; waiting will not.  Delay will allow airspeed to build, G forces on the pilot to increase (presuming a spiral or spin), and potentially the wings come off due to flutter or G forces after passing Vne -- not to mention likely exceeding the parachute's maximum deployment safety margin.
Fortunately, this kind of failure is exceedingly rare, absent the kind of damage that comes from air combat, ground impact, or extreme weather phenomena.

This one is dependent upon the control surface failure and recommended procedures from the manufacture. If the aircraft is equipped with a ballistic parachute system, the system may be activated if the pilot feels that the aircraft has departed from controlled flight with no ability on their part to regain control and recover. Typically aircraft equipped with ballistic parachutes will employ the system if a forced landing, departure from controlled flight, or other serious airborne emergencies occur.
Elevator failures have been addressed using elevator trim to control pitch and power settings.  This can give the flight crew enough time to clear in emergency and maneuver the aircraft to the nearest available airport for landing using the technique stated above to control pitch.  Aileron failures, depending on how the system fails, can sometimes be countered with roll inputs using a good aileron, if one still exists or using rudder to slip the airplane causing a shallow  roll in the direction of the rudder input.  Of all control surface failures, a rudder failure is probably the least dangerous, as the airplane still may be maneuvered using aileron and elevator inputs albeit with coordinated flight slightly compromised.  on a multi engine aircraft, yaw control may be affected by means of differential thrust using the throttles. It would be advisable in the event of a rudder failure to land the airplane at an airport with runways offering the lowest possible crosswind component on approach.
Large aircraft utilizing digital flight control computers and hydraulically actuated control surfaces may be able to detect compromised flight controls and re-route commands to existing functioning flight controls to achieve a desired response from a flight crew command.
If the aircraft is equipped with a ballistic parachute, and flight and straight and level attitude is still possible, even with the control failure, the decision to use the chute may be up to a risk management decision: Is it advisable to simply deploy the chute and ride it down, or attempt a landing with compromised flight controls.  These decisions are largely open-ended problems, and often come down to evaluating how compromised the control of the aircraft is, the pilots proficiency at handling such an emergency, the environment, and other resources available to the crew. SRM/CRM is never a closed ended problem.


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In which part of the flight does a plane reach maximum lift?
In which part of the flight does a plane reach maximum lift?
My logic:
I think a plane reach maximum lift in a horizontal flight, when a plane start rolling on a runway, landing gear produce more friction and when landing, flaps and landing gear generate more fricition with the air.
In conclusion i think the maximum lift is like this:  Horizontal flight > Takeoff > Landing

You seem to be confusing max lift to minimum drag.  If you haven't heard the term L/D Max before, (best lift to drag ratio) I suggest you research that.  It may be what you are looking for...

If it's still on the runway, it seems obvious that it's not generating as much lift as if it were climbing, no?

Do you understand the difference between lift as a force in pounds or Newtons, and lift as  a coefficient?

Part of this question seems redundant with-- https://aviation.stackexchange.com/a/56476/34686 -- key sentence "Clearly, Lift is less than Weight in a powered climb"

The above was for a linear climb. But think about what happens whenever the flight path is curving ("bending") upwards. A curvature requires a "centripetal force", which means an excess force above and beyond what is required to oppose gravity. If the flight path is initially near horizontal and then curves upward, lift will be > weight until the upward curving stops (or until the flight path has curved well above horizontal).  Isn't this a good description of what happens the instant the wheels leave the ground, and for a second or two thereafter?  )

Of course then you have banked turns which are kind of an obvious case--

The real answer is whenever the altitude is greatest :)

"Maximum lift" can mean maximum lift coefficient, maximum attainable lift at a given point in the flight envelope, or maximum attained absolute lift at any point throughout the flight envelope. Which do you mean?

In a steep turn, at high speed.
In addition to the lift needed to compensate weight, a turn requires the wing to produce additional lift to accelerate the aircraft sideways. Speed is required so enough dynamic pressure is available to produce all that lift.
The same happens at the bottom of a tight loop when weight and the lift needed to compensate centrifugal forces add up.
A third high-lift scenario happens when the airplane flies at high speed through a strong vertical updraft. In order to not overstress the structure, maximum speed is limited in gusty weather.

Would it be correct to say that the lift is proportional to whatever the G meter (if equipped) is reading?

@FredLarson: Yes.

@FredLarson in addition it is proportional to the airplane mass (as lift is force, not acceleration) too, which can be also variable even in flight (fuel consumption).

In level flight the lift must equal the weight of the aircraft (which will typically reduce as it burns fuel).  If lift exceeds weight then the aircraft will accelerate upwards; thus is something that we typically see at takeoff, and hopefully just before landing when the aircraft‘s rate of descent drops to almost zero.  The maximum lift would be generated when the aircraft pulls out of a steep dive, converting kinetic energy into lift which would allow more lift than could be generated by the power of the engines alone.


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How to calculate TAS and fuel for Diamond DA20-C1?
The PoH for the Diamond DA20-C1 is here: http://support.diamond-air.at/fileadmin/uploads/Canada/Tech_Pubs_DA20-C1/AFM/DA202-C1-Rev-28.pdf.
The plane has an MT prop, and the supplement for it is here: https://static1.squarespace.com/static/59b981e8b07869d96fe0c365/t/5e457224a5db6b6a6ce00bc2/1581609513640/DA20+MT+Propeller+Supplement+%283%29.pdf.
I need to calculate fuel consumption and TAS under the following conditions:

8000' pressure altitude (PA)
2600 RPM
-1C OAT

There's two charts in the supplement that I think I need to solve this:

It seems I need to use the first graph to find the density altitude (DA) from the temperature and PA. So for -1C and 8000, the DA appears to be 8000 as well. Looking at the table, this would give me 7.1 GPH and 132 TAS. Is this correct? I'm confused by the "standard temperature" heading in the table.

Thanks for the link, however, I don't see how I can use it to solve my problem. If you know what to do, would you mind explaining your ideas in an answer? I think I'm suppose to stick with the PoH and prop supplement to solve this.

The standard temperature line represents air in standard conditions forgiven density altitude, which is also equal to the pressure altitude


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Why are you required to commit to a full stop landing if reverse thrust is selected?
According to Airbus:

‐ After the flight crew selects reverse thrust, they must perform a full stop landing.

Does it really make sense to have this limitation, and why? What happens if you realise there's not enough space to land, and you've still got adequate speed?

The biggest problem is that thrust reversers take time to move. During that time they are still producing reverse thrust (even if only at idle) and slowing you down. They must all completely close before you get forward thrust and can add power to start accelerating again.  Then, what happens if they don't stow, or only part of them stow?  Now you've used up valuable runway that could have been used to stop (or at least slow down more) and are no longer in a position to takeoff from. The consequences can be pretty severe.

Once they're out, you're committed to stop because you don't know if, and how quickly/symmetrically, they'll stow.  Even with a long, long runway, if you push up the power and one reverser is still out, you're going off the side -- one side of the aircraft is producing forward thrust, the other is producing reverse thrust, you're done!
All the safety interlocks & features in the TR's are all about preventing uncommanded deployment -- since that might well be unsurvivable in flight.  But when you HAVE commanded the deployment, those safety features are all satisfied.  There isn't any similar level of engineering to guarantee that they stow, and do so simultaneously, because you're simply TOLD, don't try to take off once the TR's are out.
If one fails to stow at the end of the landing roll, the effect is pretty minimal with the power at idle and the aircraft at or near taxi speed, so in general there aren't great issues with the TR's perhaps still out after you're done with them.  The one case where it's a grave issue is handled with a policy solution ("don't") rather than an engineering solution.
As far as realizing that you touched down too long and you need to take it back airborne rather than trying to stop, that IS possible... right up until the point that you deploy the TR's.  Then you're committed.

Because if you do try to go around after the reversers deploy, you run the risk of them failing to retract - or, much worse, of just one of them failing to retract.
As happened on 11 February 1978, when a 737-200 flying Pacific Western Airlines Flight 314 touched down at an uncontrolled airport (Cranbrook/Canadian Rockies, CYXC) during a snowstorm, deployed its thrust reversers... and saw a snowplow further down the runway (Calgary ATC, in charge of the flight, screwed up their estimation of how long it would take to reach Cranbrook, and the runway was still in the middle of being plowed when Flight 314 landed).  During the resulting attempted go-around, the right thrust reverser restowed, but the left one did not.  The 737 cleared the snowplow, but it (the 737, not the snowplow) then, to quote the official accident report,

...climbed to 300 to 400 feet above the airfield, banked steeply to the
  left, lost height and side-slipped into the ground to the left of
  the runway.

This was despite the crew applying full right rudder and ailerons in an attempt to counteract the deployed left reverser.
42 of the 49 on board were killed as a result.

Theoretically no, you could depart again, once the thrust reversers retract and properly stow, provided you have enough runway left to takeoff and obstacle clearance at the departure ends of the runway.  If you landed on the dry lake bed at KEDW with over 15,000 ft of runway to work with, slowed down using thrust reversers and then closed them and attempted a takeoff roll, it's  possible you could again get airborne.
Of course, this generally is not the case on a revenue flight where your landing rolls will exceed several thousand feet, depending on aircraft type, local conditions, etc. you simply would not have enough runway left to attempt a second takeoff roll without a taxi back.

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/


text:
Can lift be increased by increasing the angle of incidence of the wing?
Can the lift of a plane be increased by increasing the Angle of Incidence (angle between the chord line of the main wing and the axis of the fuselage)? Meaning, can more lift be generated while the plane is in a horizontal position (i.e. not pitching nose up or down) if the wings are angled leading edge up.

Comments are not for extended discussion; this conversation has been moved to chat.

Yes.  Angle of Attack is included in the Lift formula.  Lift generally increases with angle of attack in a linear fashion until AOA reaches stall.  This is why it is not a good idea to fly at too high an AOA.  Better to increase Velocity.  Lift increases with Velocity squared.
The Navy Vought F-8 Crusader had a "variable incidence wing" that was raised to a higher pitch relative to the fuselage for take-off and landing.  Interesting to note the variable incidence wing would also improve forward visibility on landing because the fuselage nose would have a lower pitch to the horizon while the wing was at a given angle of attack.  Trailing edge flaps also raise the wing chord incidence to the fuselage with similar benefits.
While in flight, most planes simply pitch the entire plane with the elevator to increase Angle of Attack.

Yes, it is possible and variable-incidence wings have been used.
The only production example was the Vought F-8 Crusader, used primarily by the US Navy. It had a variable-incidence wing, which tilted nose-up by about 7 degrees to give increased lift for takeoff. The conventional solution of a lengthened nose undercarriage was deemed impractical for the design.
An earlier prototype was the German Blohm & Voss BV 144 transport of WWII, which was built in France and seems to have flown a few times around the end of hostilities. Variable incidence was provided because the tail was too close to the ground for conventional rotation of the whole plane at takeoff.
Another rather obvious application is the tiltwing V/STOL configuration, of which several prototypes have also flown. The additional lift provided by increased incidence can significantly ease the transitions between vertical and forward flight modes.

All it means is the deck angle of the fuselage will be lower for a given speed.  Weight being the same, the wing will fly at the AOA it needs to fly at to support the weight. By increasing the incidence, the fuse will just be pointing down more because AOA will be the same for a given flight condition.
You have a downstream problem as well because the decalage with the horizontal tail has also increased so you have to modify it (move it LE up) to maintain the same decalage angle you had before to maintain the same pitch authority and trim behaviour.

If I am thinking about this correctly, increasing angle of incidence would increase lift because you would also be increasing angle of attack. Increasing the angle of attack would increase lift until you reach the critical angle where you have airflow separation and aerodynamically stall the wing. It’s the angle of attack that’s important.
A better question is what are the advantages and disadvantages of increasing angle of attack by increasing angle of incidence. One advantage that I can think of is a reduction of drag generated by the fuselage and empennage at low speeds such as slow flight. However, there will be an increase in drag created by the wings at high speeds such as cruise flight. Intuitively, you would also assume there would be an increase in stall speed. How would an increase in incidence angle affect maneuvering such as banking for turns. Wouldn’t the increased angle of incidence exacerbate stall characteristics even when the turn is coordinated?
Just throwing some thoughts out there.

Consider what were the specific advantages motivating the design of the variable-incidence wing on the F-8 Crusader... a nice chance to exercise one's google-foo...

In short, no.  But it depends...  Angle of incidence does not increase or decrease lift, only angle of attack and airspeed affect lift.
Let me explain - If you had an aircraft with a variable angle of incidence you could fly at a constant angle of attack while varying the incidence and have no effect on lift.  Picture the wing steady with respect to the airstream, while the fuselage tilts up and down; all you would affect is drag.
Now, if the fuselage remained constant while you varied incidence then lift would be affected, but it would be affected because the angle of attack was changing!

But, if the attitude of the fuselage is constrained to stay horizontal as was stated in the question...

Then it seems the answer is yes, as per your last sentence, which possibly might better have served as the main thrust of the answer...

I'm not so sure that the no-pitch constraint was in the wording that I answered to, but whether it is or not, my point is that it is the AoA that increases lift, not incidence.  This question is really a mess...


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What does a High Pressure Turbine Clearance Control do?
In CFM56, I read that it is supposed to increase HPT performance but I don't know how it does that.

High Pressure Turbine Active Clearance Control (HPTACC) is a system that controls the temperature of the shroud of the high pressure turbine rotor (HPT). It can cool or heat the shroud support structure using bleed air from the 4th (relatively cool) or hot bleed air from the 9th stage to control the (tip) gap (or tip clearance) between the rotor (rotating part) and the shroud (stationary part).
Why take all the trouble of controlling the temperature?
Metals expand, but metal in the gas path (e.g. rotor blades) warm up much faster than the surrounding structures when power settings change. Differences in expansion from uneven heating and stresses from rotation (the higher the spool speed the higher the stress the higher the elongation), especially during transient operation (low to high or high to low power setting). This makes it very difficult (but not impossible, there are e.g. also seal types that use abradable honeycomb material in combination with knife edges) to have a single gap between the rotor and the shroud. If a single gap size would have been chosen it would be too big as it has to prevent rubbing at all times. If you can control the temperature of the shroud structure you can, based on measurements and design information, create a system that controls the gap between the shroud and the rotor blade by expanding and contracting the shroud structure.
The image below shows a typical impression of the radial position of the shroud (top orange line) of which the position is based on a displacement by thermal expansion (δth), and the tip of the HPT rotor blade (bottom displacement blue line) of which the displacement is based on thermal expansion and an elongation by rotation (δth + δm) as function of power setting. The tip clearance clearly is not constant but a function of power setting, active clearance control can be used to minimize the tip clearance by controlling the expansion of the shroud structure.

If you can keep the gap between the rotor blade and the shroud as small as possible, less air (which is a mixture of combustion products, cooling air from the high pressure stator vane and compressor discharge air) will pass over the top of the blades. The more air passes the rotor over the blade top (tip) the less will be available to extract work from to power the high pressure compressor (HPC), the less efficient your core will work.

The turbine blade clearance is the gap between the tip of a turbine rotor blade and the turbine shroud:

CFM56-7B HPTACC principle, adapted from Airbus engine systems manual
The blade is part of the rotor, and the shroud is part of the stators assembly. The length of a blade and the diameter of the shroud vary with the temperature of gas from the combustion chamber (which is on the left side above). However they don't vary exactly in the same proportion, therefore the tip clearance also varies.
When the gap is wider, a greater proportion of gas do not turn the turbine disk but escapes directly to the next stage, which in the engine shown (CFM56-7B) is the LPT since the HPT has only one rotor. This reduces the engine efficiency.
To prevent such loss the active clearance control (HPTACC) cools the turbine shroud with air from the compressor to reduces its diameter. Air is tapped at two levels in the HPC: 4th (colder) and 9th stage (hotter) and mixed to obtain the desired cooling air temperature.
This air in injected in 3 circular manifolds around the HPT (in light green above). These manifolds have many holes to diffuse air near the shroud. The shroud is cooled according to the ratio of 4th stage/9th stage air, the shroud diameter is reduced accordingly and the clearance too.
Cooling air temperature is controlled by the HPTACC valve:

CFM56-7B, HPTACC functional diagram, source Airbus engine systems manual


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Why in a CFM56-3 is vibration measured in displacement for N1 and in velocity for N2?
N1 is measured in mils (thousandth of an inch) and N2 is measured in IPS (inches per second). Why are different methods used?

Related: What does VIB stand for in EICAS?

Can you provide a source for your statement?

Sorry I should've made myself clear. Yes, the vibration sensor is located in No. 1 bearing and turbine rear frame. From CFM56-3 Engine Shop Manual in Engine Operating Limits section, the vibration level limit for N1 is in mils and for N2 is in ips.

Both seem very odd units for vibration, as it is oscillation that by definition has frequency and amplitude. Neither alone tells enough about the nature and more importantly, the energy of the vibration.

@Jpe61 both are measured synchronously. i.e. at whatever frequency the corresponding rotor is spinning.

@DanielK oh, your answer explains this quite well, thanks!

So first, both measurements are made by accelerometers, and accelerometers are by definition sensitive to acceleration.  So the original units are either G or m/s^2.  They are converted to different units for analysis and display.
So, what units are most convenient for analysis?  It depends on the frequency range.  If we have a sinusoidal vibration with displacement amplitude A and frequency $\omega$, then the corresponding velocity will be $A \omega$, and the corresponding acceleration will be $A \omega^2$.  So for very low frequencies, i.e. $\omega$ very small, so that acceleration and velocity will both be small.  Thus of the three, displacement will be the largest, and most convenient to work with.  For very high frequencies, i.e. $\omega$ very large, then acceleration i.e. $A \omega^2$ will be the largest, thus more convenient to work with.  For intermediate frequencies, velocity is the largest and thus most convenient.
N1 happens to be a fairly low frequency, due to the slow speed of the fan rotation, so displacement is good.  N2 is a medium frequency due to the faster (not but ultra fast, relatively speaking) core rotation.

Because the first sensor is measuring vibration levels on a bearing inside the engine (the #1 bearing) and the second sensor is measuring vibration at the turbine rear frame. The tolerances for vibration on those two are very different. A bearing's tolerance for vibration is extremely small, which is why it is measured in mils. The rear frame's tolerance for vibration is much larger in comparison, which is why it is measured in inches per second.

This is incorrect.  Although the sensor is mounted on the bearing, it is not primarily measuring vibration generated by the bearing itself, it is measuring vibration generated by the fan rotor.  The best place to measure rotor vibration is as close to the rotor as you can get, namely the bearing.  Similarly, the sensor on the rear frame is measuring vibration generated by the core rotor (through a stiff load path down to the bearing).  It has nothing to do with the "tolerance" of the part for vibration.

You are comparing what it is measuring vs the source of the force it is measuring. Apples and oranges. I also think you are well aware that bearings can fail. The way to tell if it is failing is to check for deviation from expected performance. If the bearing is failing, one of the primary symptoms is excessive vibration. The same for the rear frame. If you think the answer can be better worded, have at it and edit it. But the answer is correct.

Yes, excessive vibration is a symptom of bearing failure.  However, based on the comment "From CFM56-3 Engine Shop Manual in Engine Operating Limits section, the vibration level limit for N1 is in mils and for N2 is in ips" the OP is clearly asking about the synchronous fan 1/rev and core 1/rev vibration, which is very specifically related to rotor unbalance.  The choice of vibration units for these quantities has nothing to do with the tolerance of the mounting location for vibration or bearing health monitoring.

Again, I said nothing about the mounting location. The OP did.


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Does/did Russia or the Soviet Union ever fly rocket planes?
Did Russia or the Soviet Union ever have their own experimental or military rocket planes (other than the Buran), or did their military merely use jets and other non-rocket aircraft? America flew those many experimental X-planes, did the Soviet Union have any similar planes? Afaik all MiGs are jets.

Are you asking if the Soviets ever built any rocket airplane? If so it's a yes/no answer, it would make a better question if you were more specific about the uses of the rocket airplane.

@GdD I clearly wrote "experimental or military rocket planes". This may include suborbital spaceplanes similar to the X-2 and the X-15. But Mike Sowsun actually answered my question, and the question and answers in ymb1's link are interesting too.

5 minutes or less of searching would have supplied you with an answer, why didn't you bother to do that?

@jwenting You're one of the "friendlier" persons around here, right?

@Giovanni I try to be, but I like most people here, have not much time for people who don't show any attempts to solve their own problems before asking for help.

@jwenting Why did you take your time then? I didn't know what exactly to google, and I didn't know of any Russian rocket planes so far. I thought it's best to ask here to get a definite answer. It doesn't hurt you if I ask this. If you don't wanna answer, don't do it (the question had been answered anyway, your comment was entirely superfluous, only good for showing detest).

Also I'm asking specifically about experimental planes like the X-planes, not just any rocket plane.

Yes
First, we should clarify the fact that a "rocket engine" is a variant of a jet engine. A rocket engine is usually understood as a jet engine that requires external supply of all components of combustion, including the oxidizer.
In a way, that makes things easier, and this is why rocket engines were commonly used at the early stages of jet propulsion development.
Early examples include rocket boosters for traditional piston-driven airplanes. This method is still in use to shorten the take-off distance (down to zero in some cases), but early on it was tried for "regular" emergency power and, of course, research. (Ram jets also had some limited use then).
The first Soviet 100% jet was BI-1 (БИ-1) in early 1942.

wiki Commons
BI-1 had a rocket engine burning kerosene and nitric acid. Even though the airplane was designated as "fighter" and had a space provision for guns, it was really a research ariplane, with the maximum time under power of only a few minutes (it never reached even two minutes in real tests).
The first one was corroded away by the acid. The third one (BI-3) was lost, and the test pilot Bakhchivandzhi died, presumably due to Mach tuck, which was not understood then.
In the later times, during the "true" jet era, there was much less cooperation between the aviation and space industries than in the West. This is one of the reasons why "rocket planes" were a rarity in the USSR, even for research.


text:
Can we bring a rotorcraft like Ingenuity to an altitude of the same pressure as the one on Mars?
At Perseverance's and Ingenuity's location the atmospheric pressure is about 750 Pa (0.11 psi). On Earth, we find that pressure about 110,000 ft high. If we built a copy of Ingenuity, could we bring it up there or close to that altitude? Probably not, because of the Earth's higher gravity and its atmosphere's lower density, but how high could we fly it?
The highest vertical rotorcraft so far went to "only" 40,800 ft (12.4 km), this was a massive crewed helicopter.

You’re right - although the atmospheric pressure is the same at that altitude, the gravity is significantly higher and so the helicopter would not be able to fly.

I'm pretty shure it couldn't get that high on its own, starting from the ground, because of the limited battery capacity. But still it would be interesting to know how much density it would need to fly in earth's gravity.

@UlliT But let's suppose Ingenuity's copy is solar-powered.

@Giovanni Not sure adding mass in the form of solar panels would make things better. Ingenuity is very heavily optimized for low mass. They literally looked for every gram to be removed from the total. If you make such drastic changes you would basically have an entirely different vehicle altogether. I'm also not sure where you would put solar cells on a vehicle like that.

@UlliT Oh, I see.

In the first half of the video, it's not really flying, it's more like hopping up and then crashing back down. In the second half of the video, it is suspended from a cable that takes off 2/3rds of the weight in order to simulate Mars gravity.

Different but related: Could the Mars helicopter Ingenuity have been test-flown outdoors at JPL just for the heck of it? (not that they would)

I think it could, it would just have a short flying time and may require some type of launcher to get it up there...

@Giovanni I believe the record is up around 42,500 feet.  Pretty much every rotary wing altitude record is now held by a AS350.

In a past life I routinely lofted 6-12 lb (3-5kg)  payloads to the 110,000 level.  If you can design a helicopter drone that can fly at 110,000 feet, I know how to get it there inexpensively and release it.

Using the simple momentum theory, the thrust generated by a rotor is:
$ T = \sqrt[3]{2 \rho A P^2} $
where $A$ is the rotor area, $\rho$ is the density and $P$ the power needed to spin the rotor. Since Ingenuity uses two coaxial contra-rotating rotors, this value has to be doubled:
$ T = 2 \cdot \sqrt[3]{2 \rho A P^2} $
The highest thrust is normally required for hovering where it simply equals the weight:
$ T = W = mg $
where:

$m$ is the mass
$g$ is the gravity

As a proof, let's introduce on the right part of the equation the values for Ingenuity and let's see if we get the correct thrust i.e. weight on the left. So:

$\rho = 0.01 kg/m^3$ Note: density on Mars changes according to season from 0.01 to 0.02; we use the lowest value in order to be conservative
$A = \pi \cdot 0.6^2 m^2$
$P = 350W$

Considering the martian gravity of $g_{mars} = 3.72m/s^2$ this returns a mass of $1.88kg$ which matches quite perfectly the official mass of $1.8kg$. So, even on Mars, Earth's physics is right
Now, according to the question we have to bring Ingenuity on Earth at some $30km$ height and see if it can still hover. Well, the right part of the equation remains the same since at $30km$ height the density is the same as on Mars. But the left part doesn't match anymore because the gravity is now $g_{earth} = 9.5m/s^2$ i.e. some 2.5 times higher than on Mars. So no, at 30km height, gravity is too high on Earth to let Ingenuity hover.
Anyway if we go down to some $16km$ height then the density rise to $0.15kg/m^3$ giving again a perfect match between thrust and weight. Obviously other aerodynamic or structural issues have not been considered and could play an important role.


text:
How does the B737's 'throttle offline' setting work?
I'm concerned about this article:

Throttle online means that the crew can override the autothrottle logic, while Throttle offline means that the logic cannot be overridden.

So for example - rejected take-off:
Table shows that at 60 knots 'throttle mode is offline' and we cannot override.
So at 60 knots we can't first move throttle to idle (because it's blocked?) before turning off A/T? Or we can same as at 80 knots?

That article is about somebody's simulation. In the real aircraft, the pilots can always override the A/T and can always turn it off entirely. The A/T will push the throttles toward the target position, but not so hard you can't move them easily yourself. (You may need to turn them off to get them to stop trying to push the power up, but for an action like pulling them to idle for an aborted takeoff before THR HOLD engages at 84 knots, the pilot can easily override them.)
Also, the terms "online" or "offline" aren't used with reference to the authrottles in any Boeing documentation that I've seen.  They may be off, or they can be on -- in which case the mode may be "armed" or any of various other modes. But not on/off line.
Mostly the info in that article is okay, but the table referenced seems to reflect some misunderstanding about what the pilots can/can't do with the thrust levers... at least, with regard to the actual aircraft. That info may be spot-on with regard to the simulation the author is writing about.


text:
Calculating the Thrust vs RPM Relationship of a Brushless DC Motor
hope everyone is doing okay.
I am conducting a project with my friends for an Unmanned Aerial Vehicle. We are first using MATLAB - Simulink to design and simulate the quadrotor. We completed the Kinematics part of it now we are about to incorporate the rotor dynamics of the design.
We are going to be using Emax CF2822 1200kV Brushless Motors. We want to model this motor on Simulink as well. We have a general Brushless DC-Motor equation which yields RPM as its output. However, since we do not have the motors with us at the moment (we are awaiting for our funds for this project to be deposited), we are searching through the internet to find some experimental data. However, there is not too much to look at. We are going to be using 10x4.5 propellers and this setup has only the following data:
Prop: 10x4.5, Cell Count: 3S, RPM/V: 1200kV, RPM: 6100, Thrust: 745 grams, Max Current: 14.5A.
I looked for similar setups with similar motor characteristics on the internet, however I was not very successful at finding similar setups which had the potential to help the process. I am just wondering if the above data would be sufficient to have an approximate equation/relationship for the RPM vs Thrust characteristics? When I looked on the internet, I mostly saw that the relationship was non-linear about the mid-parts and kind of linear at the end parts.

I may be missing something, but thrust vs RPM would be a characteristic of the propeller, not the motor. Are you looking for power vs RPM?

@TomMcW Yes, I saw that it depends on the type of propeller chosen. However, I found that people logged Thrust data from the same propeller under different RPMs. Then, they did Curve-Fitting to relate the RPM to the Thrust for the same propeller mathematically. However, the motor I have does not have many data on Thrust values. So, I just wondered if I would be able to get a similar equation if I have just the data above?

You'd be better off working on the device to test the motor/prop for thrust/rpm so you can do the measurement when they arrive. The manufacturers don't publish the info you need.

1200 kilovolt??


text:
Starter of engine cfm56-3C
Pls ask me i really need to know this.
After the engine CFM56-3C start and light up can i disconnect it's starter from it? In order to put it on onother engine for starting process?!
Thank you in advance.

Are you asking about the starter that is connected to the accessory gearbox? You want to remove that while the engine is running?

Yes the starter that is connected to accessory gear box. I want to remove it after the engine start and stability. Can i?!

The starter air discharge duct is used by the VSV system as a fuel drain, and is connected to a bleed valve in the 5th stage of the compressor. I doubt just removing the starter is sufficient. Anyway the MEL would not allow to fly with only one engine restartable in flight. Picture of the AGB, the starter is connected to the big vertical air duct, just above the generator.

Thank you so much my friend

If your boss really wants to do this (or presses you for reasons why you can or cannot do this), you should contact CFM product support (https://www.cfmaeroengines.com/customer-support/) and ask them. They are the only ones that can truly answer a question like this.

Thank you so much

Why the downvotes? It's a valid theoretical engineering question. In which aspect is it less clear/useful than Can a fighter jet be stolen? or Is this plane landing or departing? which had 138 upvotes and only one downvote? All the contrary, upvote it!

Thanks for your support

What's suggested amounts to taking an engine apart while it is running. The amount of risk involved for the mechanics would be immense, if it's even possible at all.
If one makes enough assumptions about what risks one is willing to accept, what condition the running engine is in afterwards (presumably still running, not necessarily airworthy), how extensive the preparations are for such a stunt (i.e. somebody may have to create a procedure & checklists from scratch), and so on, all sorts of crazy things can be possible to do.
But as far as remotely practical or safe, no, taking something as big as the starter off of a running engine isn't something you're likely to see happening.
Some turboprops (the C-130, for one) can start the engine with the starter removed -- a high-speed taxi up to about 100 knots and then a gentle abort, on a sufficiently long runway is all that's needed.  But pulling (and padding!!!) the starter from a running engine? No, thanks!

@MhZerafat You want a reference for why it is a bad idea to open and take apart a running turbofan engine? How about "common sense"?

I know this is a terrible idea. I just wanna know is this possible with all the risks that can cause bad events.


text:
What is the maximum range over open water with twin-engine to require dinghy in EASA?
What is the maximum range for a twin-engine aircraft to fly over open water before having the requirement for flotation devices (vests / dinghy) under EASA-rules?


text:
Aircraft engine communication with satellite
Do pilots get an audible indication when the engines are communicating with the manufacturer whilst in flight. Do these "pings" or "hanshakes" begin on start up and are they always one hour apart?

I don't know the answer to the other questions (I think it's manufacturer and operator dependant), but nearly 100% sure that there is no audible indication.

In your question, you're mixing two things about aircraft communications:

The protocol used by the satellite transceiver aboard the plane (SDU) to communicate with the its ground station (GES). It allows for example to log on/off with the satellite service. This protocol is a custom one.

The protocol used by engines and other aircraft systems to communicate to a destination on the ground. This protocol is ACARS, it uses whatever link is currently available, VHF or satcom.

Regardless of this confusion, the crew is not aware of the data transfer associated with engine monitoring.
Satcom protocol
The "pings" you mention are part of the satellite link protocol and are not related to engines or any actual users of the satellite link aboard. Pings triggered by the ground station and sent by satellite at 60 minutes intervals to allow the satellite provider to know if a mobile customer (the aircraft) is still active and in sight of the satellite. The aircraft crew is not notified when they happen as there is no need for.
This type of interrogation is realized using a "handshake": The ground station says "Are you still there?" and the SDU answers "Of course I am". If the SDU doesn't answer then some satellite resources can be released by the ground station.
Ping existence was brought to public attention after MH-370 tragedy since the time between interrogation and reply allowed the satcom provider Iridium to compute the aircraft range, like with a radar, reducing the possible crash locations to a circular arc.
This protocol is activated as soon as the aircraft SDU contacts the satellite and as long as it doesn't "log off". In practical it is active when the aircraft has no VHF communication available. Activation is under the avionics control.
Satcom SDU can be seen as a modem. It manages the satellite link and receives/transmits data using the satellite media. Pings are part of the low-level satellite logic occurring in the background, they don't generate any notification to the crew. The crew just aware of the satellite link status (there is a FMS page reserved for it).
ACARS
The SDU protocol maintains the satcom link available to actual users. One of these users is the ACARS messaging system for which one satcom channel is reserved.
Engines use ACARS to exchange messages with stations on the ground, either at the airlines maintenance services, or at the manufacturer. This can happen any time. ACARS messages are routed through any radio link available, this can be a satellite link when flying over oceans or a traditional point to point radio link like VHF (VHF 3 transmitter is used). As satcom links are expensive, VHF links have a higher priority for ACARS.

Satcom / VHF use in flight, source
Whether ACARS applications leave a track of their activity depends on the application. For example some information can be shown to the crew on the avionics display or printed on the cockpit printer with an audio signal associated.
However, the crew is not aware of ACARS messages dedicated to engine monitoring, as again there is no need for.

This is great info, but I'm not sure it answers the question, which seems to be whether the pilots are notified of this communication. The answer is a definite no. There are no cockpit indications, although as you noted, some ACARS/satcom status is available on the CDU from the FMS (if the pilot chooses to pull up these pages).

@TypeIA: It seemed to me "pings themselves are not announced as there is no need for" answered the question. But if this is no clear enough I will update it.

I guess you did answer it, but it was buried in the details. Thanks.


text:
How does the Boeing 737's implementation of RNAV and VOR approaches compare to Airbus?
How does the Boeing 737's implementation of RNAV and VOR approaches compare to Airbus?

where have you encountered this? can you provide sources on which you base your question? are you looking for differences in pilot behaviour? or technical implementation? or something else?

Please don't delete your previous question. You could have edited that one instead of asking a new question. Please have a look at the tour to learn more about how the site works.

@federico technical implementation

An RNAV approach in the Boeing 737 is usually flown in LNAV/VNAV mode, meaning the Flight Management Computer (FMC) provides both lateral and vertical guidance. A VOR approach can be flown by manually dialing in the VOR frequency into the NAV radio and dialing in a course, which can then be tracked using VOR/LOC mode. Vertical guidance may be provided by VNAV, but the pilots can also manually dial in a vertical speed (V/S). However, the 737 can also fly a VOR approach in LNAV/VNAV mode, meaning the same way an RNAV approach is flown.
On the Airbus A300 and A310, the VOR approach works quite similar to the 737: you can manually dial in the VOR frequency and a course and follow that using NAV and V/S mode. Alternatively, the VOR approach can also be flown in profile mode, where the FMC provides lateral and vertical guidance (no need to dial in the VOR frequency any more). The same profile mode is also used for RNAV approaches, which allows flying RNAV and VOR approaches the same way.
Starting with the A320, Airbus removed the option to manually track a VOR. The only possibility to fly VOR approaches is to use managed mode (the new term Airbus uses for guidance provided by the Flight Management and Guidance Computer, FMGC). The same mode is used for RNAV approaches (if flown using LNAV/VNAV mode, not true when using LPV using an SBAS, but I don't think that's what you are asking about). All non-precision approaches (NPAs) are flown this way, so there is no difference between VOR or RNAV approaches, however the FMGC may (depending on equipment) rely on the VOR data (see related incident).
Both Boeing and Airbus offer another way to fly NPAs though (regardless of VOR or RNAV or whatever NPA approach), which is more similar to a precision approach. Boeing calls this Integrated Approach Navigation (IAN) and Airbus calls this FMS Landing System (FLS). The indications on the PFD and the procedures to fly this approach are then very similar to an ILS approach:

(image sources: Boeing 737 IAP, Airbus Safety First: FLS)


text:
What airplane is this?

I came across these images (probably taken around 2014-2015, by Arthur le Fol) of a pop group sitting on the engines of an airplane, but I don't recognize the plane. (Images are links to the sites they came from). What airplane is this? Where could it have been taken?

Oops; didn't do a good job searching this time. Turns out the images are excerped from a music video.


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How to determine the length of the CD nozzle?
I am designing a convergent nozzle as well as a CD(Converging-Diverging) nozzle. How do I find out the length of those nozzles? If the throat of the CD nozzle has 0 length then what should be the the length of converging and diverging section?

To convert as much of your built up pressure energy, your nozzle should be long enough for the built up pressure to decrease to ambient pressure or slightly below. This is true for both a convergent nozzle, and the convergent-divergent nozzle
The length of the different sections depend on the structure of around the nozzles. How much pressure can it handle, what is the maximum temperature it can handle.
You should transition from convergent to divergent, when the velocity of the air is almost supersonic.
You should then end your divergent part, when the pressure is equal to, or slightly lower than ambient pressure.

Extra stuffs
The reason you may want overexpand the air and have the output pressure be slightly lower than ambient pressure, could be to contain the air behind the engine in a straight line
It's what gives us shock diamonds in the exhaust

The exhaust stream will contract as it leaves the nozzle at ambient pressure from cooling.  This may contribute to the diamond pattern as the stream recompresses and reexpands (almost like a spring).

@RobertDiGiovanni Definitely a contributing factor!

The big factor however comes from overexpanding. For example with an exhaust pressure at 0.5 and ambient pressure at 1, the ambient pressure will naturally move towards the lower exhaust pressure.


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Why pressure and temperature decreases along the length of a convergent nozzle?
I know that the pressure and temperature follows stagnation relations of pressure and temperature as well as area mach relation inside a convergent nozzle along the length. And according to that equation, it forms a curve that represents the pressure anf temperature. But I want to know why it is happening?  Why pressure and temperature decreases along convergent nozzle?  Like what physics it follows?  What theory it follows?  What law?  According to what effect this is happening?

See How do diffusers turn velocity into pressure?, if's for a divergent nozzle, so you just need to take of the opposite reasoning. Also see this page: Pressure is inversely proportional to area. Then why is this pressure decreased in convergent nozzle.. If your question is different, then please give more details.


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How to turn inviscid data to viscous?
I have used Panel method to calculate Cp vs x/c graph and CL. But as panel method assumes attached flow it does not give an flow separation and thus in CL vs alpha graph I could not find any stall. My question is, I have all the inviscid datas, how do I turn into viscous flow data? I know that I have to involve reynold's number, skin friction coefficient and all to make it stall in CL vs alpha curve. How to I do that? What formulas, equations or method I should use?

You can still use the panel method and implement along with it an integral method for the boundary layer (like it is done in the xfoil software). The lift will be modified with the theory of displacement while for the drag the viscous drag will be computed directly by considering the friction coefficient found thanks to the integral method. One of the most common integral method (and one of the most used at my knowledge) is the Thwaites's method. You can find a reference here and an awesome review about the different methods here. I describe here one of the possible procedure:

You define the stagnation point on your profile, starting from the pressure coefficient distribution obtained from your panel method
Strating from the stagnation point the BL that is developing is laminar, hence to compute the friction coefficient (and the displacement boundary layer) you will use initially the Thwaites method. Here I will limit in describing this method. If you want to know more about transition you can couple with Michel criteria and you will find all the information in the hereby reference.

The Thwaites method is an empirical method based on:
 The Von Karman integral equation for the BL:
$$
\frac{d \theta}{dx}+\frac{\theta}{U_e}\left(2+H\right) \frac{d U_e}{dx}=\frac{1}{2}c_f
$$
 The shape factor (we define as $\delta^*$ the displacement BL thickness)
$$
H=\frac{\delta^*}{\theta}
$$
 A local Reynolds number
$$
Re_{\theta}=\frac{U_e\theta}{\nu}
$$
 The Thwaites parameter
$$
l=\frac{1}{2}Re_{\theta}c_f
$$
 The Polhausen parameter, that gives an index of the local pressure:
$$
\lambda=\frac{\theta^2}{\nu}\frac{d U_e}{dx}
$$
The Von Karman equation has three unknowns $($\theta$, $H$, $c_f$ )$, and hence we will need 2 closure equations. The first one is an empirical relation:
$$
\frac{U_e}{\nu}\frac{d\left(\theta^2\right)}{dx}=2\left[l-\left(2+H\right)\lambda\right]\approx 0.45-6\lambda
$$
By substitution we arrive at the following ODE:
$$
\frac{d}{dx}\left(U_e^6\theta^2\right)=0.45\nu U_e^5
$$
It is here that you will use your velocity computed with your panel method. in fact $U_e$ is the sum of the asymptotic velocity and of the perturbation velocity computed with your non viscous method. In this way you will obtain $\theta$ and it will be possible for you to compute the necessary quantity for each point of the discretization.
Starting from the stagnation point you will need a first $\theta$. This is approximated throught the following empirical relation:
$$
\theta=\sqrt{\frac{0.075 \nu}{U_e/R_0}}
$$
For the points that will come after this, after having computed $\theta$ you can obtain $\lambda$ using a second empirical relation and finally the friction coefficient $c_f$:
\begin{equation}
    \lambda=\frac{\theta^2}{\nu}\frac{dU_e}{dx}
\end{equation}
\begin{equation}
    l(\lambda) = \left\{ \begin{array}{rcl}
0.22+1.5\lambda-1.8\lambda^2 & ,& \lambda>0\\
0.22+1.402\lambda+\frac{0.018\lambda}{0.107+\lambda} & , & \lambda<0
\end{array}\right.
\end{equation}
\begin{equation}
    c_f=\frac{l(\lambda)}{1/2 Re_{\theta}}
\end{equation}
once obtained all the local $c_f$ it is possible to obtain the total $c_{d,f}$ by the integration of the local contribution over the profile.
Through the computation of the displacement thickness you can start to correct your lift coefficient with the viscous data.
To predict the stall the queation is a bit more complicated. I suggest you to have a look at the paper I proposed you for the transition and  to this paper. Even if it is not of the best quality is quite explicative.

How can I use panel method datas to compute velocities for viscous mode? I mean I have the slope angles for all the panels, beta angles, phi angles, length of each panels, tangential velocities of each panels, pressure coefficients of each panel, coordinates of all the control points and all. These all, I have stored in different columns of excel. How to manipulate these datas to turn into viscous mode datas? I am a bit confused how can I use these datas in Thwaites method? Please help!
datas are here--> https://drive.google.com/file/d/1DKVyhafMxI-0isfE0NcqylkEgVNyGQE4/view?usp=sharing

I tried to answer tou in a more extensive way.

Thank you so much! I can not explain how much I am relieved... Thank you so so much for helping me...! You are awesome, man

Can you give me your LinkedIn ID so that I can connect with you and get to know you more? I have calculated so many things, and things are getting so complicated at each step. I believe I am so close but still, I am not able to calculate the CL vs alpha curve accurately and get the curve to stall. We can also discuss these things in gmeet/ video call if you could just give me some of your time. I am already following you on twitter but can not start a conversation there cause twitter does not allow that. However, I am just asking for a healthy scientific discussion for a scientific research.


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What is propeller pitch speed?
I was using Ecalc to calculate some flight parameters and then I came across this term called "propeller pitch speed". Can anyone explain me what this means?

"Pitch Speed: the speed of the airflow leaving the propeller blade (nominal pitch speed based on the geometry). the theoretical maximum speed of your model neglecting aerodynamic drag of the model itself and propeller slip might be higher". This is from the user guide of this tool.

Just to add to the comment above from @mins that prop pitch speed [cm/min] = linear pitch in cm x  RPMs,

@xxavier: It seems this is another (and confusing) name for forward velocity.